(1) Field of the Invention
The present invention relates to landing gear, to an aircraft having such landing gear, and to a method implemented by the landing gear.
The invention thus lies in the technical field of landing gear, and more particularly landing gear for aircraft capable of landing vertically, and in particular rotary wing aircraft. The problems associated with airplanes that are unsuitable for landing vertically are different from those associated with aircraft that are capable of landing vertically. Under such circumstances, the term “running landing” is used to designate landing as performed by an airplane, as contrasted to “vertical landing” as can be performed by aircraft capable of landing vertically, such as helicopters.
(2) Description of Related Art
Independently of the nature of the aircraft, landing gear may comprise a plurality of undercarriages, each provided with at least one wheel in order to enable the aircraft to travel on the ground until it takes off, enabling the impact that results from a landing to be damped, and including a braking system for enabling the aircraft to be brought to rest over an acceptable distance.
The increasing weight and speed of aircraft, with the corresponding increase in vertical and horizontal levels of kinetic energy that need to be absorbed during a landing, have sometimes led to oleopneumatic shock absorbers being progressively adopted. Furthermore, braking systems are sometimes provided with hydraulic controls.
In addition, the increase in aerodynamic drag caused by landing gear in flight has led to an increasing interest in landing gear that can be retracted in flight into the fuselage of the aircraft. It can be understood that, for reasons of safety, it is appropriate to guarantee that such retractable landing gear is properly extended from its housing prior to landing.
To achieve that object, various systems provide for redundancy in the control of landing gear extension in order to mitigate malfunction of any one control.
In this context, manufacturers have devised hydraulic architectures that are simple and safe for enabling landing gear to be extended rapidly. Such an architecture has one hydraulic actuator per undercarriage, which actuator is connected by pipework to a fluid tank and to a hydraulic pump.
Conventionally, the fluid tank of a helicopter is located in the top portion of the aircraft, whereas on the contrary the landing gear is situated in the bottom portion of the aircraft. Pipework thus passes from one end of the helicopter to the other, where such a configuration maximizes the risk of leakage, the weight of the device, and complicates managing co-existence between the various hydraulic and/or electrical networks.
In addition, that architecture requires numerous sensors to be used in order to verify that the various members involved are operating properly.
In order to achieve the safety targets required by certification regulations, the retraction actuators are generally linear hydraulic actuators. Hydraulic directional control valves are connected by pipework to the linear hydraulic retraction actuators in order to request retraction or extension of the undercarriage as a function of an input order, which input order may be given mechanically or electrically.
Hydraulic retraction actuators are commonly used insofar as they provide good power per unit weight. Furthermore, such linear hydraulic retraction actuators are relatively insensitive to the seizing phenomenon, and this characteristic gives an acceptable level of safety. In particular, the risk of being faced with a linear hydraulic retraction actuator that opposes an emergency extension of an undercarriage under the effect of its own weight is practically zero.
Likewise, it is common practice to use a braking system acting via a hydraulic directional control valve that is controlled by pedals, either via a mechanical link or via a positive displacement hydraulic transmission. This leads to a problem of installing hydraulic pipework from the hydraulic generator circuit to the cockpit, and then to the location for controlling the brakes.
If braking is regulated by making use of a servo-valve that servo-controls hydraulic pressure to an electrical signal, it would appear to be much simpler to feed the servo-valve directly from the hydraulic generator circuit, and to control the pressure it delivers by a signal delivered by an electrical transmitter actuated by the pedals.
Under such circumstances, the state of the art presents undercarriages, each having a shock absorber, a hydraulic retraction actuator for retracting and extending the undercarriage into and from a wheel bay, and a hydraulic braking system.
That mainly hydraulic architecture presents the advantage of being reliable and effective. Nevertheless, it requires a large amount of pipework, pumps, fluid tanks, and numerous sensors all to be used.
In addition, for an aircraft of small size such as a helicopter, it is not unusual for pipework to be installed that goes from one end of the aircraft to the other.
It can thus be understood that such hydraulic architecture is relatively heavy and bulky. Furthermore, it can be very difficult to maintain the hydraulic architecture, e.g. in order to find a leak when the sources of a leak can in fact be numerous.
Architectures are known for heavy aircraft that make use of electrical control means for controlling hydraulic retraction actuators.
Nevertheless, in order to comply with safety requirements, provision may be made to duplicate or even to quadruplicate the control means used. The person skilled in the art then refers to “duplex” or “quadruplex” architectures.
Such an architecture has little impact on airplanes of large size in terms of weight. Nevertheless, the impact in terms of weight is unacceptable on an aircraft capable of landing vertically and presenting light or medium weight.
It should be observed that there are considerable differences between airplanes and helicopters, or more generally aircraft capable of landing vertically, and as a result the technical fields of airplanes that perform a running landing and of aircraft that land vertically are distinct.
An airplane that performs a running landing, such as an airliner, presents weight that is very great compared with a small- or medium-sized aircraft capable of landing vertically. Furthermore, an airplane that performs a running landing thus presents a forward speed on landing that is much greater than does an aircraft landing vertically.
These major differences have led to helicopter manufacturers to adapt undercarriages to their own needs that are different from those of airplane manufacturers, both in terms of structure and in terms of control. Given the major role played by hydraulics in the field of flight controls, and the mastery that has been achieved in this field, such control means, which are intrinsically available, and above all which are reliable, have been the preferred means for controlling undercarriages and brakes on aircraft that land vertically.
In the state of the art, reference may be made to document FR 2 887 516.
That document presents a communications network to which actuators are connected for maneuvering, steering, and braking
Document FR 2 946 320 describes a braking system for aircraft that has an electromechanical actuator acting on a pusher, the pusher applying a braking force against a disk.
Document US 2009/0187293 describes an architecture having a control module connected to sensors for sensing the proximity of an undercarriage and to a lever for verifying said undercarriage.
Document EP 2 107 273 describes landing gear provided with a landing gear leg carrying two wheels.
In addition, the landing gear has a primary actuator enabling the landing gear leg to be moved from a “landing gear retracted” position to a “landing gear extended” position, and vice versa.
Furthermore, the landing gear is provided with a hinged stay device and with a scissors linkage referred to as a “strut arrangement” that co-operates with a spring. That device enables the landing gear to be blocked in the “landing gear extended” position.
The landing gear has a second actuator for folding the scissors linkage in order to retract the landing gear leg by using the first actuator.
Document U.S. Pat. No. 3,224,713 is also known.